A gas turbine engine may be used to power various types of vehicles and systems. A particular type of gas turbine engine that may be used to power aircraft is a turbofan gas turbine engine. A turbofan gas turbine engine may include, for example, five major sections, a fan section, a compressor section, a combustor section, a turbine section, and an exhaust section. The fan section is positioned at the front, or “inlet” section of the engine, and includes a fan that induces air from the surrounding environment into the engine, and accelerates a fraction of this air toward the compressor section. The remaining fraction of air induced into the fan section is accelerated into and through a bypass plenum, and out the exhaust section.
The compressor section raises the pressure of the air it receives from the fan section to a relatively high level. In a multi-spool engine, the compressor section may include two or more compressors. For example, in a triple spool engine, the compressor section may include a high pressure compressor, and an intermediate compressor. The compressed air from the compressor section then enters the combustor section, where a ring of fuel nozzles injects a steady stream of fuel. The injected fuel is ignited by a burner, which significantly increases the energy of the compressed air.
The high-energy compressed air from the combustor section then flows into and through the turbine section, causing rotationally mounted turbine blades to rotate and generate energy. Specifically, high-energy compressed air impinges on turbine vanes and turbine blades, causing the turbine to rotate. The air exiting the turbine section is exhausted from the engine via the exhaust section, and the energy remaining in this exhaust air aids the thrust generated by the air flowing through the bypass plenum.
Similar to the compressor section, in a multi-spool (e.g., multi-shaft) engine the turbine section may include a plurality of turbines. For example, in a triple spool engine, the turbine section may include a high pressure turbine, an intermediate pressure turbine, and a low pressure turbine. The energy generated in each of the turbines may be used to power other portions of the engine. For example, the low pressure turbine may be used to power the fan via one spool, the intermediate turbine may be used to power the intermediate pressure turbine via another spool that is concentric to the low pressure turbine spool, and the high pressure turbine may be used to power the high pressure compressor via yet another concentric spool.
Gas turbine engines, such as the one described above, typically operate more efficiently with increasingly hotter air temperature. The maximum air temperature is typically limited by the materials used to fabricate the components of the turbine, such as the turbine blade airfoils. Thus, the airfoils are cooled using a variety of schemes, including directing some air discharged from the compressor section, and into and through cooling channels formed within the airfoils, to remove heat via convective heat transfer. At high temperatures, however, this convective heat transfer process may not sufficiently cool the airfoils, and a film cooling scheme is implemented. With the film cooling scheme, cooling air is injected onto the external surface of the airfoil via small film cooling holes that extend through the airfoil surface, and into the internal cooling channels. The merit of film cooling can be measured by a so-called “film effectiveness.”
In order to maximize the film effectiveness, the amount of cooling flow directed onto the airfoil outer surface is preferably maximized. Thus, the amount of cooling flow passing through, and thus the cross-sectional area of, the film cooling holes extending through the airfoil sidewall, is also preferably maximized. Moreover, it is preferable that the ratio of the length to diameter of each of the film cooling holes extending through the airfoil sidewall be greater than two. If the length-to-diameter ratio is greater than two) then the cooling flow will exit the film cooling holes fairly close to the upstream sidewall outer surface, which will further maximize film effectiveness. However, the thickness of the airfoil sidewall is, in many cases, small enough that other design constraints, such as the minimum distance between film cooling holes, cannot be met if these other constraints are met. Moreover, while a small relative hole angle is generally advantageous, as the hole angles relative to the airfoil surface are reduced, the inlets of the holes in a single coolant channel can interfere with one another, thereby reducing film effectiveness.
Hence, there is a need for a method of forming and locating holes in turbine blade airfoils that allows film effectiveness to be maximized, for a given airfoil geometry, and/or allow turbine operation at higher temperatures. The present invention addresses one or more of these needs.